Various spacecraft payloads, such as in scientific or classified applications, must have their spin axis positioned other than nominally orbit normal, for example towards Earth, throughout their orbit in order to operate properly. If the momentum of the spinning payload is not counterbalanced in some manner, torques, such as through thrusters must be applied continuously to precess the momentum vector around the orbit to keep the payload spin axis pointed in the desired direction. If the payload has significant inertia or spins at a high rate, the resulting torque requirements can place excessive demands upon the system in terms of weight, cost, and power (large magnetic torques, excessive thruster firings, etc.).
To eliminate the use of standard momentum control techniques to counterbalance the payload, apparatus for providing a dual spin zero momentum satellite have been developed. The dynamic equations of motion for these dual spin zero momentum satellites are well known and were disclosed in a book written by Peter C. Hughes entitled Spacecraft Attitude Dynamics published by John Wiley & Sons 1986, at pages 164-65. However, Mr. Hughes in this article did not discuss or provide any hardware implementations of these dynamic equations. In fact, the primary conclusion of his article was that the architecture did not provide any novel possibilities over those currently used for spacecraft stabilization since the spacecraft behaves qualitatively like a non-spinning body.
While some current spacecraft attitude control systems obey the equations in Hughes' article, they suffer from a variety of problems. For example, the current dual spin zero momentum systems use a reaction wheel to counterbalance the momentum of the spinning satellite. The primary disadvantages of these systems that utilize a reaction wheel counterbalance are that they are heavy, require more power than is necessary to counterbalance the satellite momentum using the approach of the present invention, and are very expensive.
For example, it is known that the reaction wheel in the prior DSP system stores about 1800 ft-lb-s of momentum, weighs approximately 180 lbs., and draws 40 W of power at steady state, and costs approximately $1.5 million. A further disadvantage of a satellite system utilizing a reaction wheel is that the wheel can also limit the rotation of the payload to unacceptably low rates. For example, the reaction wheel utilized in the current DSP spacecraft system has the largest momentum storage capacity of any reaction wheel currently available on the market. This reaction wheel stores about 1800 ft-lb-s so the allowable momentum of the spinning payload is also limited to 1800 ft-lb-s without redesign of the wheel.
SUMMARY OF THE INVENTION
The present invention relates to a method and apparatus for attitude control of a spinning spacecraft which is configured so that its spin axis attitude is not nominally inertially fixed throughout its orbit (for example pointed at Earth).
It is a further object of the present invention to provide a dual spin zero momentum system that is capable of generating an equivalent amount of momentum to counterbalance the momentum of a spinning payload that weighs less, requires less power, and costs less than prior spacecraft systems.
In accordance with the objects of the present invention, an attitude control system for providing dual spin zero momentum is provided. The spacecraft includes a first body positioned at a first end which is spinning in one direction. The spacecraft includes a second body that is positioned at a second end of the spacecraft opposite the first end. The second body is spinning in a direction opposite the direction of rotation of the first body. A bearing and power transfer assembly ("BAPTA") having a first end and a second end is disposed between the first body and the second body. The first end of the BAPTA is in communication with the first body causing it to rotate. The second end of the BAPTA is in communication with the second body causing it to rotate with respect to and in the opposite direction as the first end of the BAPTA and the first body. The first end of the BAPTA and the second end of the BAPTA share a spin axis and counterbalance the overall system momentum of the spacecraft such that the system momentum is nominally zero along the BAPTA spin axis.
In accordance with a further object of the present invention, the first body (body A) can be a payload module and the second body (body B) can be a bus module which are counterspun with respect to one another. This arrangement provides a satellite with nominally zero momentum along the spin axis of a bearing and power transfer assembly ("BAPTA") disposed between the payload module and the bus module. The BAPTA maintains a relative rotation rate of body A with respect to body b of .omega..sub.A *(1+I.sub.A /I.sub.B) , where .omega..sub.A is the angular velocity of the body A, and I.sub.A and I.sub.B are the moments of inertia of body A and body B respectively.
Other objects and features of the present invention will become apparent when viewed in light of the detailed description of the preferred embodiment when taken in conjunction with the attached drawings and appended claims.